Cooled turbine blade shroud

ABSTRACT

A process for forming a turbine blade comprises the step of forming an as-cast turbine blade having an airfoil portion and a tip shroud, wherein the forming step comprises forming at least one as-cast cooling circuit within the tip shroud.

CROSS REFERENCE TO RELATED APPLICATION(S)

The instant application is a divisional application of allowed U.S.patent application Ser. No. 12/362,724, filed Jan. 30, 2009, entitledCooled Turbine Blade Shroud.

BACKGROUND

There is described herein a turbine blade having a tip shroud withcooling circuits for use in high temperature applications.

Turbine blade tip shrouds can be used to provide a useful flowpath shape(conical flowpath outer diameter) and to minimize tip leakage flow toincrease turbine efficiency. Tip shrouds can also provide structuralbenefits by changing blade natural frequencies and mode shapes, as wellas providing frictional damping from the interaction between matingblade shroud segments. Tip shrouds can degrade in operation by creep(curling up of shroud edges) or oxidation if the shroud metaltemperature and/or stress exceed the capability of the material fromwhich the blade and the shroud are produced.

Historically, it has been difficult and expensive to provide coolingfeatures to turbine blade tip shrouds. As a result, blades with tipshrouds often have been limited to lower temperature stages of a gasturbine engine. Limitations in manufacturing capability have greatlyconstrained shroud cooling features, with existing designs eitherproviding lightweight, extensive cooling at great cost, simple coolingat reduced cost or thick, heavy designs which require very heavy bladesand rotors to support the large cooled shrouds.

Use of traditional ceramic core materials to form internal coolingpassages in blade shrouds results in air passages which are excessivelythick compared to the rest of the shroud geometry, leading to anexcessively thick and heavy blade tip and a very heavy blade/rotorstage. Failure can occur due to the high stress imparted by the heavytip shroud.

Other methods used in the past are open cavities closed withcoverplates, such as that shown in FIG. 1. The coverplates are weldedover machined cooling passages. The coverplates tend to be heavy and theoverall process of manufacture is expensive.

Another method used in the past is the fabrication of EDM coolingpassages. Such a method is shown in FIG. 2. Forming cooling passages inthis manner is expensive and has very limited, straight line passagegeometry limitations.

These prior processes for forming shrouds with cooling are expensive,create life debits due to welding, and can form heavy shrouds due toparasitic mass of a coverplate. Still other processes are slow as wellas expensive and provide limited cooling passage geometry capability.

FIGS. 3-5 show a large-size industrial engine airfoil concept that usesa large plenum core in the tip shroud fed by drilled holes in the blade.The dashed outline shown in FIG. 5 illustrates the plenum boundary. Theairfoil is fabricated using covers and ceramic core inserts. Thisfabrication concept suffers from being expensive and heavy. Further,this concept used a plenum, rather than a defined duct with a confinedpath with inlets and exits. Plenums such as this suffer from uncertainlocal internal flow conditions with low heat transfer.

SUMMARY OF THE DISCLOSURE

In accordance with the present disclosure, there is provided a shroudhaving a plurality of cooling passages, which cooling passages areformed using refractory metal core technology. Cooling passages formedin this manner are advantageous because they provide controlled internalair velocity and effective cooling through the extent of the passage.

A turbine blade for use in high temperature applications is disclosed,which turbine blade broadly comprises an as-cast airfoil portion and anas-cast outer tip shroud portion, the outer tip shroud portion having atleast one as-cast internal cooling passage for cooling the outer tipshroud, and the at least one as-cast internal cooling passage having oneor more exits for discharging cooling air over exterior surfaces of theshroud.

A process for forming a turbine blade is disclosed which broadlycomprises the steps of forming an as-cast turbine blade having anairfoil portion and a tip shroud, and the forming step comprisingforming at least one as-cast cooling passage within the tip shroud.

Other details of the RMC cooled turbine blade shroud of the presentdisclosure, as well as objects and advantages attendant thereto, are setforth in the following detailed description and the accompanyingdrawings wherein like reference numerals depict like elements.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates an approach for providing a cooled tip shroud;

FIG. 2 illustrates another approach for providing a cooled tip shroud;

FIGS. 3-5 illustrate a plenum approach for providing a cooled tipshroud;

FIG. 6 shows a shroud with a cross section of an airfoil superimposedthereon;

FIG. 7 is a sectional view of a shroud having internal cooling passagesformed using refractory metal core technology; and

FIG. 8 illustrates a ceramic core with refractory metal cores attachedthereto.

FIG. 9 illustrates a single refractory metal core used to form more thanone passage.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)

As described herein, there is disclosed a turbine blade having a tipshroud with a plurality of thin cooling passages cast integrally intothe tip shroud using refractory metal core technology. The passages mayhave a thickness in the range of from 0.010 to 0.060 inches. This typeof thin, as cast, internal cooling passage in the tip shroud provideshigh heat transfer with a very small increase in shroud thickness,namely from 0.030-0.100 inches less thickness than required byconventional ceramic core casting techniques.

This type of manufacturing is useful because the shape of the refractorymetal core(s) can be tailored as needed to the specific blade beingdesigned without the need for expensive machining operations and/orwelded coverplates. Heat transfer augmentation features, such as tripstrips and pedestals, can be easily fabricated and used as needed toincrease shroud cooling and passage flow.

Referring now to FIGS. 6 and 7, there is shown a turbine blade 10 havingan airfoil portion 12 and an outer tip shroud 14. The tip shroud 14 maybe provided with a first cooling passage 16 and a second cooling passage18. Each of the cooling passages 16 and 18 is formed using refractorymetal core technology. Each of the cooling passages has an inlet 20which communicates with a source (not shown) of cooling fluid via acommon central channel or fluid conduit 19 within the airfoil portion12. Each cooling circuit 16 and 18 may be desirably located at amid-plane level of the as-cast shroud. By “mid-plane”, it is meant thatthere is an equal thickness of the shroud above and below each coolingcircuit 16 and 18. Offset cooling passages may be advantageous to somespecific designs.

Each of the cooling passages 16 and 18 may have a one or more exits forflowing cooling fluid over desired portions of the tip shroud 14, suchas over exterior surfaces of the shroud, or directly out of the shroud.As can be seen from FIG. 7, the cooling passage 16 may have an exit 22on one side of the tip shroud 14 and a plurality of exits 24 and 26 onan opposite side of the tip shroud 14. The cooling passage 18 may havean exit 28 on one side of the tip shroud 14 and three cooling exits 30,32, and 34 on an opposite side of the tip shroud 14. The number ofcooling exits and their locations in each cooling passage 16 or 18 maybe tailored as needed to promote efficient cooling of the shroud. A tipshroud 14 having as-cast cooling passages 16 and 18 with the exits asshown in FIG. 7 provides efficient cooling at low cost and weight.

The turbine blade 10 with the airfoil portion 12 and the tip shroud 14may be formed using any suitable casting technique in which a primaryceramic core 100 (such as that shown in FIG. 8) is used to form theprimary blade radial inner passages with the primary ceramic core 100being centrally positioned within a die having the shape of the outerportions of the turbine blade. As can be seen from FIG. 8, a pluralityof refractory metal cores (RMCS) 102 and 104 are joined to the primaryceramic core 100. The refractory metal cores 102 and 104 may be formedfrom any suitable refractory material known in the art, such asmolybdenum or a molybdenum alloy. Each of the refractory metal cores 102and 104 may be joined to the primary ceramic core 100 by means of one ormore tabs 108 bent over and inserted into slots 110 in the tip 112 ofthe primary ceramic core 100. The turbine blade 10 with the outer tipshroud 14 may be formed by casting any suitable superalloy material in aknown manner. After the molten superalloy material has been poured intoa mold (not shown) and cooled to solidify and form the turbine blade 10,the airfoil portion 12 and the tip shroud 14, the primary ceramic core100 may be removed using any suitable leaching technique known in theart. Thereafter, the refractory metal cores 102 and 104 may be removedusing any suitable leaching technique known in the art. Once therefractory metal cores 102 and 104 are removed, there is left an as-castshroud having the as-cast, thin cooling passages 16 and 18.

If desired, the refractory metal cores 102 and 104 may each be providedwith a plurality of slots or holes for forming a plurality of pedestalsor a plurality of trip strips in each cooling circuit 16 and 18 forenhancing cooling effectiveness.

If desired, as shown in FIG. 9, a single refractory metal core 122 maybe used to form more than one passage in the finished part. The portions124 and 126 are outside the envelope of the finished casting and areremoved after the pot is formed. Also, it may be desirable to have onecooling passage, rather than multiple passages.

One advantage to the approach described herein is that the exits for thecooling circuits may be sized to provide a desirable level of coolingwithout the need to employ machining of the as-cast material. Thus, thetechnique described herein is a cost effective technique for introducingextensive cooling features in a turbine blade tip shroud, with minimalincrease in shroud thickness. This allows turbine tip shrouds to be aneffective option in engine environments where the gas temperature issubstantially above the useful temperature capability of the airfoilalloy where they were previously not practical and/or cost effective.This is of potential value for low pressure turbine blades that canbenefit from a conical OD flowpath and reduced tip leakage provided byshrouded stages.

It is apparent that there has been provided in accordance with theinstant disclosure a RMC cooled turbine blade shroud. While the RMCcooled turbine blade shroud has been described in the context ofspecific embodiments thereof, other unforeseen variations, alternatives,and modifications may become apparent to those skilled in the art havingread the foregoing description. Accordingly, it is intended to embracethose alternatives, variations, and modifications as fall within thebroad scope of the appended claims.

1-6. (canceled)
 7. A process for forming a turbine blade comprising thesteps of: forming an as-cast turbine blade having an airfoil portion anda tip shroud: and said forming step comprising forming at least oneas-cast cooling circuit within said tip shroud.
 8. The process accordingto claim 7, wherein said airfoil portion forming step comprises using aprimary ceramic core to form at least one radial inner passage withinsaid airfoil portion of said turbine blade; and said at least oneas-cast cooling circuit forming step comprises attaching a plurality ofrefractory metal cores to said primary ceramic core.
 9. The processaccording to claim 8, wherein said attaching step comprises joining eachof said refractory cores to said primary ceramic core by inserting aplurality of tabs on each said refractory metal core into a plurality ofslots in a tip of the primary ceramic core.
 10. The process of claim 7,wherein said at least one cooling circuit forming step comprises formingat least one cooling circuit at a mid-plane level of the as-cast shroud.11. The process of claim 7, wherein said at least one cooling circuitforming step comprises forming two cooling circuits at a mid-plane levelof the as-cast shroud.
 12. The process of claim 7, wherein said formingstep comprises forming each said cooling circuit with a first coolingfluid exit on one side of the shroud and at least two additional coolingfluid exits on a second side of the shroud.